System and method for vaporizing space debris in space

ABSTRACT

A system and method for vaporizing space debris in space. The system includes a spacecraft body, a primary solar concentrator mounted to the spacecraft body that collects and focuses solar flux from the sun, and a secondary solar concentrator positioned at a focal point of the primary solar concentrator that refocuses the focused solar flux. A manipulator arm coupled to the spacecraft body grabs the space debris in space and positions it at a location where the refocused solar flux vaporizes the debris. The secondary solar concentrator can be a point-source concentrator, the primary solar concentrator can be a parabolic mirror, a Fresnel lens or a light focusing element or assembly, and the space debris can be a retired spacecraft or launch vehicle upper stage or component.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a divisional application of U.S. patent applicationSer. No. 15/280,448 filed Sep. 29, 2016, titled “On-Orbit ThermalExtraction of Raw Materials from Space Debris in Support of AdditiveManufacturing of New Space Elements On-Orbit.”

BACKGROUND Field

This invention relates generally to a system and method for vaporizingspace debris in space and, more particularly, to a system and method forvaporizing space debris in space, where the system included a primarysolar concentrator for focusing solar flux from the sun, a secondarysolar concentrator for refocusing the focused solar flux and amanipulator arm for positioning the space debris at a location in spacewhere the refocused solar flux vaporizes the debris.

Discussion

Many satellites have been and continue to be launched into Earth orbitfor various applications, both military and civilian, such as forcommunications, Earth observation, scientific study purposes, etc. Thesesatellites are usually positioned in a geosynchronous Earth orbit (GEO),medium Earth orbit (MEO) or low Earth orbit (LEO). As is known in theart, a satellite in GEO has an altitude and speed that allows it toorbit the Earth at one revolution per day, thus causing the satellite toappear to remain stationary above a particular point on the Earth. Otherorbits are also available for satellites.

The launch of satellites from the Earth creates space debris as a resultof discarded rocket boosters and other components required to positionthe satellite in the desired orbit. Further, satellites that become wornout, obsolete, defective, etc., may be taken out of service and remainon orbit, also contributing to the space debris. Satellites in GEO thatare taken out of service are often moved from GEO to a higher altitude,such as 200 miles above GEO, to a graveyard orbit before they are takenout of service and become non-functioning so as to open up GEO slots forother newer functioning satellites. Also, space debris can collide witheach other creating exponentially more individual pieces of spacedebris. Thus, space debris comes in all sizes from very small flecks tofull-sized satellites and rocket boosters.

Typically, space debris orbits the Earth at a very high speed, and thuscan be a hazard to functioning and operating spacecraft and satellitesif they collide. The U.S. Government and others track many thousands ofspace debris elements, including very small components, and provide anavenue through which the position of active spacecraft and satellitescan be altered to avoid collisions. However, the amount of space debriscontinues to increase, and eventually some form of debris removalprocess will be necessary.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is an illustration of a spacecraft system including a spacecraftthat collects and refines material from space debris for manufacturingnew spacecraft systems;

FIG. 2 is a close-up illustration of a solar refinery portion of thespacecraft system shown in FIG. 1;

FIG. 3 is an illustration of a debris collector and new spacecraftfabrication and assembly spacecraft that employs solar arrays to collectsolar energy instead of the solar concentrator shown in the spacecraftsystem in FIG. 1;

FIG. 4 is an illustration of an assembly line like layout of an on-orbitspacecraft manufacturing module that uses 3-D printing, and subtractiveor vapor deposition manufacturing approaches to produce spacecraft inorbit; and

FIG. 5 is an illustration of a modification to the solar concentratingdebris conversion spacecraft that refocuses the Sun's energy from theprimary concentrator in order to achieve even higher temperaturessufficient to completely vaporize any space debris.

DETAILED DESCRIPTION OF THE EMBODIMENTS

The following discussion of the embodiments of the invention directed toa system and method for vaporizing space debris in space is merelyexemplary in nature, and is in no way intended to limit the invention orits applications or uses.

Orbital space debris can be a resource of materials since the debrisincludes the elements that are of the type necessary to make new spacesystems, such as new satellites. In order to accomplish this, it isnecessary to be able to collect the debris, separate the elements, andthen use the elements in a manufacturing system that is itself part of asatellite orbiting the Earth. The development of 3-D printing, alsoknown as additive manufacturing (AM), provides one way to depositmaterial into new useful systems. As is known, 3-D printing is a processwhere a material is fed into a heated nozzle and is laid down layer bylayer to build a desired a product. Conventional subtractivemanufacturing as well as vapor deposition and material removal processesused to produce integrated circuits on the Earth's surface are alsoprocesses that can be used to produce spacecraft on-orbit, once usablematerial has been refined, collected and stored in-orbit.

Spacecraft produced in-situ or in-orbit have an advantage overspacecraft produced terrestrially in a satellite factory becausespacecraft produced in-situ do not need to withstand the high forcesrequired for launch, including accelerating to ˜8 km/sec in less 15minutes. Spacecraft that are manufactured on-orbit, using techniquessuch as 3-D printing do not require the strong structure needed byspacecraft produced on Earth to survive launch. Spacecraft producedon-orbit can be simple flat panels, for example, with antenna elementson the face of the panel aimed downward towards Earth, solar arrays onthe backside of the panel aimed towards the sun and the various elementsneeded for a spacecraft inside the panel. These spacecraft elements canbe arranged in layers inside the panel, similar to the make-up ofprinted circuit boards made on Earth. Thus, if a viable system could beemployed to separate space debris into its various elements and thenlaid-down or assembled as a spacecraft using 3-D printing, subtractiveor vapor deposition techniques, then what is now the current problem ofthousands of spacecraft in the GEO graveyard orbit would be convertedinto a high value resource since this material can be used to producenew spacecraft and does not need the high cost of being launched intoGEO orbit where it will be used, since it is already there. Theexistence of a technology to seamlessly separate existing spacecraftmaterial into useable stock piles for on-orbit production would be atechnology worth potentially billions of dollars because it coulddeliver billions of dollars of mission value through use as new types ofcommunications and ground observation satellites. Ultimately thecombination between material separation and precision on-orbitproduction would mean that these systems can be reprocessed over andover, reducing or ultimately eliminating the need for new systems to belaunched. Small elements that are difficult to produce on-orbitinitially, such as computer processors, can always be launched fromEarth in a single launch vehicle and used to economically enable theproduction of many spacecraft in-orbit.

As will be discussed below, the present invention proposes a spacecraftsystem that is placed on orbit around the Earth, and employs elementsthat allow the system to collect space debris, such as non-functioningsatellites, selectively heating the collected debris at precisedifferent temperatures to melt the different elements in the debris atdifferent times so that they can be separately collected, and thenretrieving the separated elements as needed in an on-orbit manufacturingprocess to fabricate new spacecraft systems having updated and desiredfeatures, where it is subsequently put into operation.

FIG. 1 is an illustration of a spacecraft system 10 including aspacecraft 12 having a spacecraft body 14 and a solar concentrator 16mounted thereto. The solar concentrator 16 can be any suitable solarconcentrator for the purposes described herein, such as a parabolicreflector, Fresnel lens, light focusing element, etc., and needs to beof the proper size to generate enough focused energy to melt spacecraftelements as will be discussed below. In one embodiment, the solarconcentrator 16 is a parabolic dish 34 having a reflective surface 36and being about 100 feet in diameter. The spacecraft body 14 includesall of the elements necessary to operate the spacecraft 12 as discussedherein, including antennas, communications systems, thrusters, etc. Thespacecraft 12 also includes a solar refinery 20 having a curved crucible22, a collection element 24, a rotary actuator 26 and a manipulator 28having a number of arm sections 30 including a processing section 32,where the manipulator 28 is coupled to the spacecraft body 14.

FIG. 2 is an illustration of the solar refinery 20 shown in detailwithout the rest of the spacecraft 12 shown, but showing the crucible 22being in an open position so as to allow it to capture spacecraftdebris, such as a launch vehicle upper stage or component, representedhere as a retired spacecraft 40, where the spacecraft debris can be anydebris orbiting the Earth. It is noted that the discussion herein ofemploying the crucible 22 to collect and refine the retired spacecraft40 is by way of a non-limiting example in that any system or devicesuitable to heat and separate constituent elements based on temperaturemay be able to be employed. The spacecraft 12 is moved to a suitablelocation near the retired spacecraft 40 using onboard thrusters (notshown), or the debris can be moved to the spacecraft by another vehicle,and then the manipulator 28 is actuated to position the crucible 22directly adjacent to the retired spacecraft 40 to capture it. Once theretired spacecraft 40 is positioned within the crucible 22, the crucible22 is closed so as to confine the retired spacecraft 40 therein. Oncethe retired spacecraft 40 is secured within the crucible 22, then themanipulator 28 moves the crucible 22 relative to a focal point of thesolar concentrator 16, as shown in FIG. 1, to heat the debris.

The concentrator 16 is then oriented towards the sun 50 and solar flux52 received by the concentrator 16 is focused onto the crucible 22 as afocused beam 54 to heat the retired spacecraft 40 therein and melt aselect set of the constituent elements. The crucible 22 can be made ofany suitable high temperature material, such as graphite or tungsten,that has a high enough melting temperature, so that the heat necessaryto melt the spacecraft elements is not hot enough to heat or damage thecrucible 22. The rotary actuator 26 rotates the crucible 22 so that thefocused beam 54 is not focused on a single spot on the crucible 22 so asto help prevent the crucible 22 from being damaged or melted.Alternately, the actuator 26 can be a linear actuator. Careful movementof the refinery 20 relative to the solar flux 54 and rotation of therefinery 20 allows control of the refinery's temperature and thetemperature of its contents. The entire spacecraft 12 can also be movedfrom slightly off-pointed from the Sun 50 to be directly aimed at theSun 50 to adjust the temperature of the debris inside the crucible 22.The retired spacecraft 40 is held inside the crucible 22 so that itrotates with the crucible 22, which causes the particular element thatis being melted to be flung out by centripetal force or other forcesinduced by movement of the crucible 22 and its contents towards theinside surface of the outer walls of the crucible 22.

The manipulator 28 can position the crucible 22 in the focused beam 54so that the internal temperature of the crucible 22 is preciselycontrolled so as to sequentially melt one material of the retiredspacecraft 40 after another. As the constituent elements are heatedwithin the crucible 22 and the crucible 22 is rotated or moved, thosemelted elements are collected by collection devices (not shown) in thecollection element 24. For example, if the manipulator arm 28 positionsthe crucible 22 at a certain location in the focused beam 54 so that thecrucible 22 heats up to a particular temperature, the element in theretired spacecraft 40 that melts at that temperature will melt and becollected by a certain collection device, such as a closable chamber, inthe element 24, where that chamber is closed once the crucible 22 isheated for some period of time. An angled or curved shape to thecrucible wall can direct the molten material into the collection element24 that collects the material for use during the production phase. Themanipulator 28 then moves the crucible 22 to a different location in thefocused beam 54 so that the temperature of the crucible 22 increases anda different element in the retired spacecraft 40 is melted independentof the other elements, and is collected in a different collectioncompartment in the collection element 24 in the same manner. Thisprocess continues until all of the constituent elements in the retiredspacecraft 40 are melted and collected, where any material left over inthe crucible 22 can be mechanically collected and compressed for use asballast or other various applications. The collected material can bereheated for 3-D printing production using the solar flux 52 or usingelectrical power produced by solar arrays (not shown in FIG. 1).Material can also be vaporized and then condensed on cold plates (notshown) as a way of increasing the purity of the material, beforeeventually being melted and collected for on-orbit manufacturing usingthe same system, with cold plates added.

As discussed above, the solar concentrator 16 is used to focus the solarflux 52 to generate the heat. However, in an alternate embodiment, it ispossible to generate heat using the solar flux 52 in a different manner.FIG. 3 is an illustration of a spacecraft 60 that can replace thespacecraft 12, where like elements are identified by the same referencenumber. In the spacecraft 60, the solar concentrator 16 is replaced withsolar arrays 62 and 64 and the spacecraft body 14 includes a spacecraftfabrication module 66. The solar arrays 62 and 64 convert the solar flux52 to electrical energy in a known manner. The electrical energygenerated by the solar arrays 62 and 64 is transferred by suitableelectrical wiring (not shown) through the fabrication module 66 to, forexample, resistive heaters (not shown) provided within the crucible 22.Although the spacecraft 60 that employs the solar arrays 62 and 64 maybe more efficient for generating the heat to separate the constituentelements from the space debris, solar arrays are typically veryexpensive, and it may be more cost effective to use the solarconcentrator 16.

Once all of the elements are collected in the collection element 24,which can contain the collected material from multiple retiredspacecraft and other debris, the present invention proposes using thosenow separated elements that may be in various levels of purity tofabricate a new spacecraft system using in this example one or more 3-Dprinters, as mentioned above. The 3-D printer can be any suitable 3-Dprinter that is configurable in the manner discussed herein. The 3-Dprinter or manufacturing system can be located at any suitable locationon the spacecraft 12 or 60, for example, in the section 32 of themanipulator 28 or in the fabrication module 66. Alternately, themanufacturing system can be flying in formation with the spacecraft 12.

FIG. 4 is an illustration of a spacecraft fabrication and assemblymodule 70 generally illustrating a 3-D printing process, where themodule 70 fabricates the new spacecraft, referred to herein as apanelsat 72, as it moves or is moved along a flat fabrication backplane74 from left to right in this illustration. Alternately, otherthree-dimensional layered or pocketed structures or components can beproduced. Also, the 3-D printing or other production elements can bemoved to work on a stationary panelsat as it is built-up. This examplepanelsat 72 is defined as very similar to a single large electroniccircuit board, where the payload and spacecraft systems are integratedtogether and laid down in layers or in adjacent spaces to form acomplete spacecraft. Spacecraft sub-assemblies, boxes, wire harnesses orconventional single function bolt-on units are not required since thein-orbit spacecraft can be organically built-up with integratedstructures, harnessing and active elements as a single structure. Use ofmodular elements with different functionality is possible but notrequired. It is noted that fabrication of the panelsat 72 is by way of anon-limiting embodiment in that other configurations for otherspacecraft are within the scope of the present invention.

The module 70 includes a plurality of 3-D printers, conventionalsubtractive fabrication devices or vapor deposition and removal deviceseach performing a different printing, fabricating, layering or removaloperation in a certain sequence. In this example, the module 70 includesa structure printer 78, an antenna printer 80, an insulator printer 82,an interconnects printer 84, a propulsion printer 86, a thermal printer88 and a solar array printer 90 are used. Each of the printers removesthe material collected in the collection element 24 necessary for theprinting operation using any suitable technique, where the material maybe in molten, solid or vapor form. This fabrication process allows thepanelsat 72 to have radiators and antenna elements on one side of thepanelsat 72 that will face the Earth, all the various sub-systems,control modules, computers, etc., in the middle of the panelsat 72 and asolar array on an opposite surface of the panelsat 72 facing away fromthe Earth. The first elements could lay down a grid of active antennaelements covered by insulation material and linking structures with viaswhere power connections run in layers between each of the vias and otheractive elements. The structure printer 78 can print a relativelynon-robust structure that does not need to survive spacecraft launch,and may be a honeycomb or webbed configuration where openings in thestructure allow various subsequently printed elements to be supportedand integrated therein.

Some of the elements that may go into the panelsat 72 may not be able tobe printed on orbit as described, such as high level electronics,computers, etc. One or more high level elements, such as a plug-in box,can be brought from the Earth to provide a central computer, precisionACS elements and even propulsion functions until the technology toproduce these elements on-orbit is available. Those elements,represented as element 94 on the panelsat 72, can be plugged into thepanelsat 72 by a core electronics installer 92. The ultimate goal wouldbe to fabricate the entire panelsat 72 in space.

It is noted that although the discussion herein talks about employing3-D printing to fabricate new spacecraft systems, other manufacturingprocesses, such as conventional subtractive, vapor deposition and laseretching, may also be employed on orbit to use the constituent elementsto generate the new spacecraft system. Further, it may be possible totake advantage of the vacuum environment in space when fabricatingcertain integrated circuit chips that are fabricated by, for example,chemical vapor deposition processes.

It is further noted that the process of melting constituent elements fora 3-D fabrication printer as discussed herein can also be employed forasteroid mining, where the spacecraft 12 collects the material from anasteroid instead of space debris for subsequent spacecraft fabrication.

The discussion above talks about melting constituent elements in spacedebris that are separately collected and then used in an additivemanufacturing process to produce a new spacecraft system, where the heatnecessary to melt the elements is provided by focusing sun light orcollecting sun light by a solar array. Instead of melting theconstituent elements and then reusing them in a manufacturing process,the present invention also proposes eliminating space debris byvaporizing the debris, where the vapor is then harmlessly disbursed inspace. For this embodiment, it is not necessary to separately controlthe temperature that the elements are heated, but it is necessary tosignificantly heat the elements above the melting temperature of theelement to vaporize it.

FIG. 5 is an illustration of a spacecraft system 100 showing thisembodiment of the invention, where like elements to the system 10 areidentified by the same reference number. In this embodiment, themanipulator arm 28 is used to grab the retired spacecraft 40 andposition it at the appropriate high heat intensity location. Because thesun is not a point source, the focal point of the solar flux 52 from thesun 50 off of the concentrator 16 is not a well defined point. In otherwords, the size of the sun blurs its focused energy so that the highertemperature elements and compounds in the retired spacecraft 40 will notvaporize. Therefore, in order to obtain the heat necessary to vaporizeall of the elements on the retired spacecraft 40, such as tungsten andcarbon, the solar collector 16 would need to be uneconomically large.Thus, the system 100 includes a secondary solar point-sourceconcentrator 102 that is positioned at the focal point of theconcentrator 16 and focuses the beam 54 to be a more focused beam 104 toachieve the desired vaporization temperature. The vapor that isdispersed from the retired spacecraft 40 under the high heat isrepresented by dotted line 106. The orientation between theconcentrators 16 and 102 is shown here by way of a non-limiting examplein that this orientation can be any suitable orientation for a specificapplication.

As will be well understood by those skilled in the art, the several andvarious steps and processes discussed herein to describe the inventionmay be referring to operations performed by a computer, a processor orother electronic calculating device that manipulate and/or transformdata using electrical phenomenon. Those computers and electronic devicesmay employ various volatile and/or non-volatile memories includingnon-transitory computer-readable medium with an executable programstored thereon including various code or executable instructions able tobe performed by the computer or processor, where the memory and/orcomputer-readable medium may include all forms and types of memory andother computer-readable media.

The foregoing discussion discloses and describes merely exemplaryembodiments of the present invention. One skilled in the art willreadily recognize from such discussion and from the accompanyingdrawings and claims that various changes, modifications and variationscan be made therein without departing from the spirit and scope of theinvention as defined in the following claims.

What is claimed is:
 1. A system for vaporizing space debris in space,said system comprising: a spacecraft body; a solar collection devicemounted to the spacecraft body for collecting solar flux from the sunthat is converted into heat; and a manipulator device for grabbing thespace debris and positioning it at a location in space where the heatfrom the solar flux vaporizes the space debris, wherein the solarcollection device includes a first solar concentrator that is positionedto receive the solar flux and a second solar concentrator that ispositioned at a focal point of the first solar concentrator, said secondconcentrator being operable to receive focused solar flux from the firstsolar concentrator and refocus the focused solar flux on the spacedebris.
 2. The system according to claim 1 wherein the first solarconcentrator is a parabolic mirror, a Fresnel lens or a light focusingelement or assembly.
 3. The system according to claim 1 wherein thesecond solar concentrator is a point-source concentrator.
 4. The systemaccording to claim 1 wherein the manipulator device is manipulator armcoupled to the spacecraft body.
 5. The system according to claim 1wherein the space debris is a retired spacecraft or launch vehicle upperstage or component.
 6. A system for vaporizing space debris in space,said system comprising: a spacecraft body; a primary solar concentratormounted to the spacecraft body and being configured to collect and focussolar flux from the sun; a secondary solar concentrator positioned at afocal point of the primary solar concentrator and being configured torefocus the focused solar flux from the primary solar concentrator; anda manipulator arm coupled to the spacecraft body for grabbing the spacedebris and being configured to position it at a location in space wherethe refocused solar flux vaporizes the space debris.
 7. The systemaccording to claim 6 wherein the secondary solar concentrator is apoint-source concentrator.
 8. The system according to claim 6 whereinthe primary solar concentrator is selected from the group consisting ofa parabolic mirror, a Fresnel lens and a light focusing element orassembly.
 9. The system according to claim 6 wherein the space debris isa retired spacecraft or launch vehicle upper stage or component.
 10. Amethod for vaporizing space debris in space, said method comprising:providing a primary solar concentrator on a spacecraft body; collectingand focusing solar flux from the sun by the primary solar collector;providing a secondary solar concentrator at a focal point of the primarysolar concentrator; refocusing the focused solar flux by the secondarysolar concentrator; and positioning the space debris in space at alocation where the refocused solar flux vaporizes the space debris. 11.The method according to claim 10 wherein the secondary solarconcentrator is a point-source concentrator.
 12. The method according toclaim 10 wherein the primary solar concentrator is selected from thegroup consisting of a parabolic mirror, a Fresnel lens and a lightfocusing element or assembly.
 13. The method according to claim 10wherein the space debris is a retired spacecraft or launch vehicle upperstage or component.